Processes for producing components containing ceramic-based and metallic materials

ABSTRACT

Processes for producing a component containing a ceramic-based material and having detailed features formed from materials other than ceramic materials. Such a process entails producing the component to include a first subcomponent and at least a second subcomponent having at least one off-axis geometric feature that results in the second subcomponent having a more complex geometry than the first subcomponent. The first subcomponent is formed of a ceramic-based material, and the second subcomponent and its off-axis geometric feature are separately formed of a metallic material and attached to the first subcomponent to yield a robust mechanical attachment. The component may be, for example, a gas turbine airfoil component.

BACKGROUND OF THE INVENTION

The present invention generally relates to ceramic-based articles andprocesses for their production. More particularly, this invention isdirected to processes of producing ceramic-based articles to includemetallic regions that define detailed features, for example, dovetails,shanks, platform features and tip shrouds of gas turbine airfoilcomponents.

Higher operating temperatures for gas turbines are continuously soughtin order to increase their efficiency. Though significant advances inhigh temperature capabilities have been achieved through formulation ofiron, nickel and cobalt-base superalloys, alternative materials havebeen investigated. Ceramic materials are a notable example because theirhigh temperature capabilities can significantly reduce cooling airrequirements. As used herein, ceramic-based materials encompasshomogeneous ceramic materials as well as ceramic matrix composite (CMC)materials. CMC materials generally comprise a ceramic fiberreinforcement material embedded in a ceramic matrix material. Thereinforcement material may be discontinuous short fibers dispersed inthe matrix material or continuous fibers or fiber bundles orientedwithin the matrix material. The reinforcement material serves as theload-bearing constituent of the CMC in the event of a matrix crack. Inturn, the ceramic matrix protects the reinforcement material, maintainsthe orientation of its fibers, and serves to dissipate loads to thereinforcement material. Silicon-based composites, such as siliconcarbide (SiC) as the matrix and/or reinforcement material, are ofparticular interest to high-temperature applications, for example,high-temperature components of gas turbines including aircraft gasturbine engines and land-based gas turbine engines used in thepower-generating industry.

Continuous fiber reinforced ceramic composites (CFCC) are a type of CMCthat offers light weight, high strength, and high stiffness for avariety of high temperature load-bearing applications, includingshrouds, combustor liners, vanes (nozzles), blades (buckets), and otherhigh-temperature components of gas turbines. A notable example of a CFCChas been developed by the General Electric Company under the nameHiPerComp®, and contains continuous silicon carbide fibers in a matrixof silicon carbide and elemental silicon or a silicon alloy. SiC fibershave also been used as a reinforcement material for a variety of otherceramic matrix materials, including titanium carbide (TiC), siliconnitride (Si₃N₄), and alumina (Al₂O₃).

Examples of CMC materials and particularly SiC/Si—SiC (fiber/matrix)CFCC materials and processes are disclosed in U.S. Pat. Nos. 5,015,540,5,330,854, 5,336,350, 5,628,938, 6,024,898, 6,258,737, 6,403,158, and6,503,441, and U.S. Patent Application Publication No. 2004/0067316. Onesuch process is known as “prepreg” melt-infiltration (MI), which ingeneral terms entails the fabrication of CMCs using multiple prepreglayers, each in the form of a tape-like structure comprising the desiredreinforcement material and a precursor of the CMC matrix material, aswell as one or more binders and typically carbon or a carbon source. Theprepreg must undergo processing (including firing) to convert theprecursor to the desired ceramic. Prepregs for CFCC materials frequentlycomprise a two-dimensional fiber array comprising a single layer ofunidirectionally-aligned tows impregnated with a matrix precursor tocreate a generally two-dimensional laminate.

For purposes of discussion, a low pressure turbine (LPT) blade 10 of agas turbine engine is represented in FIG. 1. The blade 10 is an exampleof a component that can be produced from a ceramic-based material,including CMC materials. The blade 10 is generally represented as beingof a known type and adapted for mounting to a disk or rotor (not shown)within the turbine section of an aircraft gas turbine engine. For thisreason, the blade 10 is represented as including a dovetail 12 foranchoring the blade 10 to a turbine disk by interlocking with acomplementary dovetail slot formed in the circumference of the disk. Asrepresented in FIG. 1, the interlocking features comprise protrusionsreferred to as tangs 14 that engage recesses defined by the dovetailslot. The blade 10 is further shown as having a platform 16 thatseparates an airfoil 18 from a shank 20 on which the dovetail 12 isdefined. The blade 10 may be further equipped with a blade tip shroud(not shown) which, in combination with tip shrouds of adjacent bladeswithin the same stage, defines a band around the blades that is capableof reducing blade vibrations and improving airflow characteristics. Byincorporating a seal tooth, blade tip shrouds are further capable ofincreasing the efficiency of the turbine by reducing combustion gasleakage between the blade tips and a shroud surrounding the blade tips.

Because they are directly subjected to hot combustion gases duringoperation of the engine, the airfoil 18, platform 16 and tip shroud havevery demanding material requirements. The platform 16 and blade tipshroud (if present) are further critical regions of a turbine blade inthat they create the inner and outer flowpath surfaces for the hot gaspath within the turbine section. In addition, the platform 16 creates aseal to prevent mixing of the hot combustion gases with lowertemperature gases to which the shank 20, its dovetail 12 and the turbinedisk are exposed, and the blade tip shroud is subjected to creep due tohigh strain loads and wear interactions between its seal tooth (ifpresent) and the shroud surrounding the blade tips. The dovetail 12 isalso a critical region in that it is subjected to wear and high loadsresulting from its engagement with a dovetail slot and the highcentrifugal loading generated by the blade 10.

Current state-of-the-art approaches for fabricating ceramic-basedturbine blades have involved integrating the platform 16, dovetail 12,airfoil 18 and tip shroud (if present) as one piece during themanufacturing process, much like conventional investment castingtechniques currently used to make metallic blades. However, the platform16, dovetail 12, tangs 14 and tip shroud represent detailed geometricfeatures of the blade 10 that pose substantial challenges to designing,manufacturing and integrating CMC components into an affordable,producible design for turbine applications. For example, the process ofintegrating a platform 16 and tip shroud with the airfoil 18 using CMCmaterials creates complexities in the design and manufacturing process,and can result in a process that can be too expensive to be economicallypractical. Furthermore, the platform 16, dovetail 12 and its tangs 14have interface/support functions that can require structural interfacecapabilities that can be difficult to achieve with CMC materials. Inaddition, the low strain-to-failure capabilities of typical CMCmaterials and the possibility of undesirable wear interactions betweentip shroud seal teeth and conventional shrouding materials poseadditional challenges to implementing CMC materials in shrouded bladedesigns.

BRIEF DESCRIPTION OF THE INVENTION

The present invention provides processes for producing hybrid componentscontaining ceramic material, in which detailed features of thecomponents are formed of materials other than ceramic materials, yetresult in a robust mechanical attachment of the ceramic and non-ceramicportions of the components.

According to a first aspect of the invention, a process produces acomponent comprising a first subcomponent and at least a secondsubcomponent having at least one off-axis geometric feature that resultsin the second subcomponent having a more complex geometry than the firstsubcomponent. The process includes producing the first subcomponent of aceramic-based material, and then separately forming the secondsubcomponent and the off-axis geometric feature thereof and attachingthe second subcomponent to the first subcomponent. The secondsubcomponent is formed of at least one metallic material.

In particular embodiments of the invention, the component may be anairfoil component of a gas turbine, for example, a blade or vane, andthe first subcomponent of the component may include an airfoil and a nubfor retaining the second subcomponent on the airfoil. Furthermore, theone or more geometric features may include portions of a dovetail,shank, platform feature and/or tip shroud of the airfoil component.

Another aspect of the invention is a component produced by the processdescribed above, a nonlimiting example of which is an airfoil componentof a gas turbine.

A technical effect of this invention is the ability to produce certainportions of a component from a ceramic-based material, while producingother portions of the component having intricate geometric details frommaterials that do not require the temperature capability ofceramic-based materials. The invention is particularly beneficial forapplications in which the intricate geometric details formed of thenon-ceramic material are interface/supporting features that requirestructural interface capabilities, and as a result of being fabricatedfrom a non-ceramic material are not nearly as labor intensive or requirethe level of skilled labor that would be required if the entirecomponent were fabricated from a ceramic-based material.

Other aspects and advantages of this invention will be betterappreciated from the following detailed description.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a perspective view schematically representing a turbine bladeof a type formed of CMC materials in accordance with the prior art.

FIG. 2 is a perspective view schematically representing a turbine bladehaving an airfoil portion formed of a CMC material and platform anddovetail portions formed of a metallic material in accordance with anembodiment of the present invention.

FIG. 3 is a side view schematically representing the turbine blade ofFIG. 2, and showing the airfoil portion as having an integral shank nubwithin a shank portion of the blade that comprises the platform anddovetail portions in accordance with an embodiment of the presentinvention.

FIGS. 4 and 5 are isolated perspective views of the integral airfoilportion and shank nub and the integral platform and dovetail portions ofthe turbine blade of FIGS. 2 and 3.

FIG. 6 is a perspective view showing a cross-section of the integralplatform and dovetail portions of the turbine blade of FIGS. 2 and 3.

FIG. 7 is a more detailed cross-sectional view of the dovetail portionrepresented in FIG. 6.

FIGS. 8, 9 and 10 are detailed cross-sectional views showing dovetailportions in accordance with other embodiments of the present invention.

FIGS. 11, 12 and 13 are isolated perspective views of the shank nub ofFIG. 3 modified to have slots, holes and protuberances, respectively,for promoting the attachment of the integral platform and dovetailportions to the integral airfoil portion and shank nub.

FIG. 14 schematically represents a co-casting process for simultaneouslyforming and attaching the integral platform and dovetail portions on theintegral airfoil portion and shank nub of the blade of FIGS. 2 and 3.

FIG. 15 schematically represents a portion of an interface regionbetween the integral airfoil portion and shank nub and the integralplatform and dovetail portions of the blade of FIGS. 2 and 3, andrepresents a coating system at the interface for inhibiting chemicalinteractions therebetween.

FIG. 16 is a perspective view schematically representing a turbine bladehaving an airfoil portion formed of a CMC material and platform,dovetail and tip shroud portions formed of metallic materials inaccordance with another embodiment of the present invention.

FIG. 17 is an isolated perspective view of the turbine blade of FIG. 17,showing the airfoil portion as having integral shank and blade tip nubs.

FIG. 18 is a more detailed cross-sectional view of the blade tip nubrepresented in FIG. 17.

FIGS. 19, 20 and 21 are detailed cross-sectional views showing blade tipnubs in accordance with other embodiments of the present invention.

FIG. 22 is a perspective view schematically representing a turbine vanehaving airfoil portions formed of a CMC material and platform portionsformed of a metallic material in accordance with another embodiment ofthe present invention.

FIG. 23 is a side view schematically representing the turbine vane ofFIG. 22, and showing each airfoil portion as having integral shank nubsreceived within pockets of the platform portions.

FIGS. 24 and 25 are isolated perspective views of one of the shank nubsof FIG. 23 modified to have slots and holes, respectively, for promotingthe attachment of the platform portions to the airfoil portions.

DETAILED DESCRIPTION OF THE INVENTION

The present invention will be described in terms of processes forproducing components containing ceramic-based materials, includinghomogeneous ceramic materials and CMC materials that may containdiscontinuous and/or continuous fiber reinforcement materials. Whilevarious applications are foreseeable and possible, applications ofparticular interest include are high temperature applications, forexample, components of gas turbines, including land-based and aircraftgas turbine engines. Furthermore, specific reference will be made toairfoil components, including turbine blades and vanes for use withinthe turbine sections of a gas turbine engine. While the invention isapplicable to a wide variety of ceramic-based materials, ceramic-basedmaterials of particular interest to the invention are believed to be CMCmaterials containing silicon, such as CMC's containing silicon carbideas the reinforcement and/or matrix material, for example, continuoussilicon carbide fibers in a matrix of silicon carbide. However, otherceramic-based materials are also within the scope of the invention.

FIGS. 2 and 3 represent a low pressure turbine (LPT) blade 30 of a typeused in an aircraft gas turbine engine. Similar to the prior art blade10 of FIG. 1, the blade 30 represented in FIGS. 2 and 3 is adapted formounting to a disk or rotor (not shown) within the turbine section of agas turbine engine. For this reason, the blade 30 is represented asincluding a dovetail portion 32 for anchoring the blade 30 to a turbinedisk. The dovetail portion 32 is configured to interlock with acomplementary dovetail slot formed in the circumference of the disk. Asrepresented in FIGS. 2 and 3, the interlocking features compriseoppositely-disposed tangs 34 that protrude from the dovetail portion 32for engagement with recesses defined by the disk dovetail slot. Theblade 30 is further shown as having a platform portion 36 that separatesan airfoil portion 38 from a shank portion 40 on which the dovetailportion 32 is defined. Depending on its particular application and therotor disk (not shown) on which the blade 30 is to be assembled, theblade 30 may comprise additional features, for example, angelwings 42 onits shank portion 40 and a shroud at its blade tip (for example, asrepresented in FIGS. 16 and 17).

Similar to what was described for the blade 10 of FIG. 1, the airfoilportion 38 and platform portion 36 are directly exposed to hotcombustion gases during operation of the engine, and the platformportion 36 is a critical region of the blade 30 in that it creates theinner flowpath surface of the hot gas path for the hot combustion gases,and creates a seal to prevent mixing of the combustion gases with lowertemperature gases internal to the rotating system and to which the shankportion 40, its dovetail portion 32 and the turbine disk are exposed. Inaddition, the dovetail portion 32 is subjected to wear and high loads asa result of its engagement with the disk dovetail slot and the highcentripetal loading generated by the blade 30.

The airfoil portion 38 of the blade 30 is an excellent candidate forbeing produced from a ceramic-based material, and especially a CMCmaterial, because it is directly exposed to the hot combustion gases andhas a generally linear geometry. On the other hand, the platform portion36, dovetail portion 32 and its tangs 34 have more complex geometriesthan the airfoil portion 38, in the sense that the airfoil portion 38has a generally linear geometry along its dominant axis, whereas thedovetail and platform portions 32 and 36 define geometric featuresoriented transverse to each of their dominant axes. Furthermore, theseoff-axis geometric features are detailed interface/supporting featuresof the blade 30, and therefore require structural interface capabilitiesthat pose substantial challenges to designing, manufacturing andintegrating a completely CMC blade (such as the blade 10 of FIG. 1) intoan affordable, producible design for turbine applications. The presentinvention provides a process for taking advantage of thehigh-temperature capabilities of CMC materials, while avoiding thedifficulties of producing complicated geometries from CMC materials. Inparticular, the present invention involves producing the airfoil portion38 and a nub 48 of the shank portion 40 as a unitary piece from a CMCmaterial, and producing one or both of the platform portion 36 anddovetail portion 32 from materials other than the CMC material used toproduce the unitary airfoil portion 38 and shank nub 48.

As used herein, the term shank nub refers to a limited portion,preferably an interior region, of the entire shank portion 40, whichfurther includes the dovetail portion 32 and its tangs 34. Asrepresented in FIG. 3, the shank nub 48 is entirely encased in thematerial used to form the platform portion 36 and dovetail portion 32.In addition, the shank nub 48 can be referred to as being “defeatured,”in that the detailed dovetail features conventionally required for ashank (such as the dovetail 12 and tangs 14 of the shank 20 of FIG. 1)can be completely omitted from the shank nub 48 shown in FIGS. 2 and 3.

As a ceramic-based material, the unitary airfoil portion 38 and shanknub 48 can be produced by known ceramic processes. For example, theunitary airfoil portion 38 and shank nub 48 can be CMC materialsfabricated from prepregs. Nonlimiting examples include the processesdisclosed in U.S. Pat. Nos. 5,015,540, 5,330,854, 5,336,350, 5,628,938,6,024,898, 6,258,737, 6,403,158, and 6,503,441, and U.S. PatentApplication Publication No. 2004/0067316. As a particular example, theunitary airfoil portion 38 and shank nub 48 can be fabricated by thepreviously-described prepreg melt-infiltration (MI) process, whereinmultiple prepregs are formed to contain the desired reinforcementmaterial and a precursor of the CMC matrix material, as well as one ormore binders and, depending on the particular desired CMC material,possibly carbon or a carbon source. The prepregs undergo lay-up, aredebulked and cured while subjected to elevated pressures andtemperatures, and subjected to any other suitable processing steps toform a laminate preform. Thereafter, the laminate preform may be heatedin a vacuum or an inert atmosphere to decompose the binders and producea porous preform that is then melt infiltrated. If the CMC materialcontains a silicon carbide reinforcement material in a ceramic matrix ofsilicon carbide (a SiC/SiC CMC material), molten silicon is typicallyused to infiltrate into the porosity, react with a carbon constituent(carbon, carbon source, or carbon char) within the matrix to formsilicon carbide, and fill the porosity. However, it will be apparentfrom the following discussion that the invention also applies to othertypes and combinations of CMC materials.

Because of the generally linear geometry of the airfoil portion 38 andshank nub 48, the required lay-up process is not nearly as laborintensive and does not require the level of skilled labor that would berequired if the entire blade 30 were to be fabricated from prepregs.FIG. 4 represents an example of a unitary CMC subcomponent 44 thatcomprises the airfoil portion 38 and shank nub 48, is entirely formed ofa CMC material, and can be produced by a CMC process such as thatdescribed above. As represented, the shank nub 48 comprises an enlargedknob or base 49 that is wider in cross-section than the region of thenub 48 adjacent the root of the airfoil portion 38. This uncomplicatedfeature can also be formed with the CMC process, and is desirable forassisting in the retention of the dovetail portion 32 of the blade 30.

Though the drive for additional turbine engine performance has promptedthe desire for using CMC materials due to increased gas pathtemperatures, those regions of blades (and other turbine components)that are not directly exposed to the hot combustion gases, including thedovetail, platform and shank portions 32, 36 and 40 of the blade 30, mayutilize materials with lower temperature capabilities, for example,nickel-, cobalt- or iron-based alloys currently available and used inturbomachinery. Notable but nonlimiting examples include suchsuperalloys as IN (Inconel) 718, René N5 (U.S. Pat. No. 6,074,602), RenéN6 (U.S. Pat. No. 5,455,120), GTD-444®, René 77 (U.S. Pat. No.3,457,066), René 80, René 80H and René 125. FIG. 5 represents a unitarysubcomponent 46 that, in combination with the CMC subcomponent 44 ofFIG. 4, preferably yields the complete blade 30 of FIGS. 2 and 3.According to a preferred aspect of the invention, the subcomponent 46 isformed of one or more of the aforementioned metal alloy materials. Theutilization of a superalloy to form the dovetail, platform and shankportions 32, 36 and 40 of the blade 30 addresses numerous producibilitylimitations that exist with current state of the art CMC processes, andalso allows for the use of known lifing methodologies and analyticaltools to verify suitable designs for the blade 30 and particularly thedetailed interface/supporting features of the blade 30, which in FIGS. 2and 3 are represented by the platform portion 36, dovetail portion 32and its tangs 34.

As evident from the CMC subcomponent 44 seen in FIG. 4, the subcomponent46 of FIG. 5 is not configured to be prefabricated and then assembledwith the CMC subcomponent 44 by inserting the shank nub 48 into acomplimentary cavity 52 defined in the subcomponent 46, though such anapproach is not outside the scope of the invention. Instead, oneapproach that has been developed during investigations leading to thisinvention involves forming the subcomponent 46 by casting metallicmaterial around the shank nub 48 of the CMC subcomponent 44. Thisapproach is practical in view of typical CMC materials having higherprocessing temperatures than the casting temperatures of a wide range ofmetallic materials suitable for forming the metallic subcomponent 46,which therefore allows the merging of established metallic castingprocesses with the CMC subcomponent 44. As a result, the prior necessityto fabricate the entire blade 30 from a CMC material is avoided, as arethe difficulties encountered when trying to produce intricate shankgeometric details in a production environment, as well as thedifficulties encountered when attempting to analyze and correlatein-service operational conditions of an all-CMC blade. Instead, all ofthe detailed features of the blade 30, and particularly the detailsassociated with its dovetail, platform and shank portions 32, 36 and 40,can be produced by machining the as-cast metallic material usingexisting manufacturing techniques.

In view of the above, a metallic material can be cast around the shanknub 48 of the simplified, de-featured CMC subcomponent 44, to producethe entire unitary metallic subcomponent 46, which in effect is anoverlaying metal casing that defines the dovetail portion 32 and itstangs 34, as well as what will be referred to as a shank casing 50 thatencases the portion of the shank nub 48 above the dovetail portion 32 ofthe subcomponent 46. In FIGS. 2, 3 and 5, the unitary metallicsubcomponent 46 is further represented as defining the platform portion36, such that the shank casing 50 is disposed between the dovetail andplatform portions 32 and 36. Additional features can also be defined bythe metallic subcomponent 46, including the angelwings 42 extending fromthe shank casing 50.

Because the coefficient of thermal expansion (CTE) of metallic materialsthat can be used to form the cast metallic dovetail portion 36 and shankcasing 50 is typically higher than the CTE of typical CMC materials,during solidification of the metallic material around the CMC shank nub48 the cast metallic material that defines the cavity 52 will contractmore than the CMC material and compress the shank nub 48, providing acompression fit and tight encapsulation of the CMC nub 48 and retentionof the CMC subcomponent 44 to the metallic subcomponent 46, which is inaddition to the retention capability provided as a result of thesubcomponent 46 surrounding the enlarged base 49 of the shank nub 48. Asa nonlimiting example, a suitable compression fit is believed to beachievable with a metallic material such as the aforementionednickel-based superalloy René 80H, which has a CTE of about 14 ppm/° C.,in comparison to a CTE of about 4 ppm/° C. for SiC—SiC CMC materials.This CTE differential is capable of yielding a strain of about 1% whencooled to room temperature from a casting temperature of about 2200° F.(about 1200° C.), resulting in a room temperature stress state in whichthe CMC shank nub 48 is in compression and the metallic subcomponent 46surrounding the nub 48 is in tension. Particularly for blades whosedovetails are in compression during operation, the shrink-fit resultingfrom the casting process is capable of providing a robust mechanicalattachment.

FIG. 6 is a perspective view showing a cross-section of the blade 30 inthe region of the interface between its subcomponents 44 and 46,evidencing in more detail the manner in which the metallic subcomponent46 can be used to completely encase the shank nub 48, including its base49. FIG. 7 is a detailed end view of the dovetail portion 32 representedin FIG. 6, evidencing how the metallic subcomponent 46 fully encases thebase 49 of the shank nub 48, and in doing so defines pressure faces 35that will engage surfaces of the disk dovetail slot in which the blade30 is to be installed. FIG. 8 is a similar end view showing a bladedovetail portion 32A installed in a disk dovetail slot 54 of a turbinerotor disk 56. The dovetail portion 32A is similar to the dovetailportion 32 of FIGS. 2-7, but differs in that the metallic material hasnot been cast to cover the lower surface of the shank nub base 49.Instead, the subcomponent 46 covers the shank nub base 49 to the extentnecessary to define the pressure faces 35. FIG. 9 is a view similar tothe view depicted in FIG. 7, but depicts a dovetail portion 32B havingmultiple sets of tangs 34B and 34C instead of the single set of tangs 34depicted in FIGS. 2-8. Furthermore, the metallic material has not beencast to cover any part of the shank nub base 49. Instead, thesubcomponent 46 covers surfaces of the shank nub 48 above its base 49 sothat the metallic subcomponent 46 is cast to define the tangs 34B andthe pressure faces 35 of the dovetail portion 32 in their entirety. Thebase 49, in effect, defines the lower set of tangs 34C in theirentirety, which are subjected to lower loads (if any) due to thereliance of the pressure faces 35 defined by the upper set of tangs 34C.Finally, FIG. 10 is similar to FIG. 9, but depicts metallic material ashaving been cast to cover the entire shank nub base 49, with the resultthat the metallic subcomponent 46 defines the dovetail portion 32B andboth sets of tangs 34B and 34C in their entirety.

FIGS. 11 and 12 are isolated perspective views of the shank nub 48 ofFIG. 3 modified to have slots 60 and holes 62, respectively, forpromoting the attachment of the CMC subcomponent 44 to the metallicsubcomponent 46. The slots 60 are defined as limited recesses in theshank nub base 49, whereas the holes 62 preferably pass entirely throughthe shank nub 48 above its base 49. In each case, metallic materialenters the slots 60 and/or holes 62 during the casting process, suchthat solidification of the casting material creates complementaryinterlocking metallic features (not shown) within the slots 60 and/orholes 62. In the case of the holes 62, the casting material within theholes 62 is also capable of interconnecting those portions of the shankcasing 50 separated by the shank nub 48. The interlocking effectphysically promotes the retention capability provided by the shank nubbase 49, and therefore further promotes a robust mechanical attachmentbetween the subcomponents 44 and 46. Though FIGS. 11 and 12 show theslots 60 and holes 62 as alternative configurations, combinations ofslots 60 and holes 62 are also within the scope of the invention.Furthermore, other negative surface features (depressions or recesses)could be defined in the shank nub 48 and/or its base 49 to achieve asimilar effect. As depicted in FIG. 13, positive surface features(protuberances) 63 can also be defined in the shank nub 48 and/or itsbase 49 and, alone or in combination with recesses (such as slots 60and/or holes 62) employed to achieve a retention effect similar tonegative surface features.

The process of “co-casting” the metallic subcomponent 46 on the CMCsubcomponent 44 can be achieved in a variety of ways. FIG. 14schematically represents a co-casting process for simultaneously formingand attaching the metallic subcomponent 46 and its unitary dovetail,platform and shank portions 32, 36 and 40 on the CMC subcomponent 44 andits unitary airfoil portion 38 and shank nub 48. As represented in FIG.14, the process may be performed within a mold 64 to produce a casting66 whose shape approximates the final geometry desired for the dovetail,platform and shank portions 32, 36 and 40, and thereby minimize theamount of post-cast machining required of the casting 66 to produce thesubcomponent 46. The mold 64 may be any suitable design, for example, aceramic shell. Following solidification of the metallic material, themold 64 can be removed to retrieve the co-cast subcomponent 46 that hasbeen cast in-situ onto the shank nub 48 of the CMC subcomponent 44.Notably, this casting technique is preferably performed to avoid thehigh temperatures and long time exposures normally required inconventional investment casting processes, which could result inundesirable chemical reactions between the CMC material of thesubcomponent 44 and the molten metal material (for example, theformation of silicides), as well as with the mold 64. The process oframping the shank nub 48 of the CMC component 44 into the moltenmetallic material within the mold 64 enables the contact time andtemperature to be kept to a minimum to prevent undesirable reactions.

Other methods that can be used to form the metallic subcomponent 46include spin casting techniques. As known in the art, spin castingprocesses are similar to conventional investment casting processes inthe fact that a mold is created by coating a wax replica of the part inceramic, and then removing the wax to yield a female form the part(“mold”), which is then filled with molten metal that solidifies to formthe final part. Spin casting techniques depart from conventional castingmethods in that the latter relies on gravitational force to act on amolten metal to fill the mold, whereas the mold in a spin castingprocess is rotated to induce centrifugal forces that act on the moltenmetal. This additional force is beneficial to certain casting geometriesand/or materials to ensure a complete fill of the mold with acceptablemicrostructure and lack of internal defects. Spin casting also differsfrom centrifugal casting processes, in which a molten metal is pouredfrom a crucible into a central pour cup that is aligned with therotational axis of a rotating mold. The molten metal initially has zerocentrifugal force acting upon it, and takes a finite amount of timeuntil it flows away from the center of rotation and slowly picks upcentrifugal force. With spin casting, the charge (unmelted raw material)is melted at a distance way from the center of rotation, such that whenthe charge is melted and rotation starts, the molten metal isimmediately acted upon by centrifugal force, resulting in a more rapidfill of the mold than either conventional or centrifugal castingprocesses.

It should be noted here that the subcomponent 46 depicted in FIG. 5could be prefabricated and then assembled with the CMC subcomponent 44.For example, the subcomponent 46 can be fabricated as two or more piecesthat can be assembled around the shank nub 48 and then welded or brazedto each other to form the complete subcomponent 46. However, thisapproach would require precision machining to control interface contactstresses between the CMC and metallic subcomponents 44 and 46 andachieve an effective level of compression and encapsulation of the CMCshank nub 48 comparable to that possible with casting techniques.However, an advantage to this approach is the ability to use alloys withmelting temperatures that would be otherwise incompatible with the CMCmaterial of the subcomponent 44, for example, due to posing an excessiverisk of reactivity or exceeding the temperature capability of the CMCmaterial. With a prefabrication technique, it may be possible to fillgaps between the CMC subcomponent 44 and the individual pieces of themetallic subcomponent 46 during assembly of the pieces. For example,gaps could be filled during the assembly process with a powdered brazefiller material, which is then sufficiently melted during brazing tojoin the pieces of the subcomponent 46 together. Brazing temperatures,for example, in a range of about 2200 to about 2300° F. (about 1200 toabout 1260° C.), would be compatible with most CMC materials currentlybeing contemplated for the invention.

In addition to or as an alternative to relying on the casting techniqueto minimize undesirable chemical reactions between the CMC material ofthe subcomponent 44 and the molten metal material of the subcomponent46, the present invention also contemplates the use of interfacecoatings provided between the CMC and metallic subcomponents 44 and 46.In addition or alternatively, an interface coating can be employed toenhance thermal expansion compliance for the shrinkage of the metalaround the CMC subcomponent 44 during solidification to reduce theincidence of cold cracking. FIG. 15 schematically represents a portionof an interface region between the CMC and metallic subcomponents 44 and46 of the blade of FIGS. 2 and 3, and represents a coating system 70 atthe interface for inhibiting chemical interactions therebetween. Thecoating system 70 can be produced to have any number of coating layersformed of a variety of different materials, and can be deposited withthe use of a variety of processes, including slurry coating, air plasmaspraying (APS), and so forth. In FIG. 15, the coating system 70 isrepresented as comprising two distinct layers 72 and 74, though thecoating system 70 could be formed by a single layer or more than twolayers, as an example, five layers. The layer 72 directly contacting theCMC subcomponent 44 may be formed by, for example, a material that isparticularly compatible with the material of the CMC subcomponent 44,for example, a glass and/or liquid-phase forming material. Examples ofpotential glass materials for the layer 72 include materials describedfor a reaction barrier coating disclosed in U.S. patent application Ser.No. 12/984,836 to Shyh-Chin Huang et al., whose contents regarding thereaction barrier coating are incorporated herein by reference. The layer74 directly contacting the metallic subcomponent 46 may be formed by,for example, a material that provides a suitable transition between thelayer 72 and the metallic subcomponent 46 in terms of chemical andphysical compatibility. For example, the layer 74 may have a gradedcomposition in which its composition immediately adjacent the layer 72is the same as or otherwise compatible with the glass and/orliquid-phase forming material of the layer 72, while its compositionimmediately adjacent the subcomponent 46 is the same as or otherwisecompatible with the metallic material used to form the subcomponent 46.For example, the layer 74 or at least its composition immediatelyadjacent the subcomponent 46 may contain or consist of a ductile metalfoam material that is chemically compatible with the metallic materialof the subcomponent 46 and provides thermal expansion compliance betweenthe CMC and metallic subcomponents 44 and 46. Suitable materials for thefoam material are believed to include high-temperatureoxidation-resistant alloys such as iron-, cobalt- and nickel-basedalloys, notable but nonlimiting examples of which include FeCrAlY alloysof types known in the art. In combination, the layers 72 and 74preferably confer a degree of compliance to the coating system 70,enabling the coating system 70 to serve as a compliant interface thataccommodates shrinkage of the metal subcomponent 46 around the CMCsubcomponent 44 during solidification. Generally speaking, thicknessesof about 0.005 to about 0.040 inch (about 0.1 to 1 millimeter) arebelieved to be suitable for the coating system 70, though lesser andgreater thicknesses are also foreseeable.

FIGS. 16 and 17 represent an LPT turbine blade 80 that is shrouded,whereas the blade 30 in FIGS. 2 through 15 is unshrouded. The blade 80comprises a CMC subcomponent 82 that defines an airfoil portion 84. Theblade further comprises a dovetail portion 86 and a platform portion 88.Optionally, the CMC subcomponent 82 may be formed to have a shank nub(not shown) surrounded by a metallic subcomponent that defines thedovetail and platform portions 86 and 88, generally in a manner similarto that described for the blade 30. Contrary to the blade 30, the blade80 represented in FIGS. 16 and 17 comprises a metallic subcomponent 90that defines a shroud portion 92 at the tip of the CMC airfoil portion84, which effectively defines a nub 94 for attachment of the shroudportion 92 to the airfoil portion 84. The metallic subcomponent 90 isfurther represented as defining an integrated seal tooth 96. Themetallic subcomponent 90 can be formed in the same manner as thatdescribed for the metallic subcomponent 46 of the blade 30.

As evident from FIGS. 17 and 18, the blade tip nub 94 can have a shapesimilar to the shank nub 48 depicted in FIGS. 4 and 6-13, for the blade30. In particular, the blade tip nub 94 is wider in cross-section thanthe immediately adjacent region of the airfoil portion 84, and serves toassist in the retention of the metallic subcomponent 90 and its shroudportion 92 at the tip of the CMC airfoil portion 84. Also similar to theshank nub 48, the blade tip nub 94 can further incorporate positive andnegative surface features that, during the co-casting process to producethe metallic subcomponent 90 and its shroud portion 92, result in thecreation of integral complementary interlocking metallic features. Asrepresented in FIGS. 19, 20 and 21, the nub 94 can have through-holes98A (FIGS. 19 and 20) and/or slots 98B (FIG. 21) of various sizes andshapes. In the case of the through-holes 98A, metallic material entersthe holes 98A during the casting process, such that solidification ofthe casting material creates complementary interlocking metalliccrossbars that extend entirely through the holes 98A to interconnectportions of the shroud portion 92 separated by the nub 94. In the caseof the slots 98B, the casting material within the slots 98B createscomplementary interlocking metallic ribs that extend entirely throughthe slots 98B to interconnect portions of the shroud portion 92separated by the nub 94. The interlocking effect physically promotes theretention capability provided by the nub 94, and therefore furtherpromotes a more robust mechanical attachment between the CMC andmetallic subcomponents 82 and 90. Though FIGS. 19 through 21 representonly negative surface features, positive surface features(protuberances) could be defined in the nub 94 as an alternative or inaddition to the holes 98A and/or slots 98B to achieve a retention effectsimilar to negative surface features.

FIGS. 22 and 23 represent a gas turbine vane segment 100 as anothersuitable application for the present invention. The vane segment 100 isrepresented as having two airfoil portions (vanes) 102 between a pair ofinner and outer platforms (bands) 104, though a single airfoil portion102 or more than two airfoil portions 102 could be present. The vanesegment 100 is one of a number of vane segments that are assembledtogether to form an annular-shaped vane assembly of a turbine engine.The airfoil portions 102 are excellent candidates for being producedfrom ceramic-based materials because they are directly exposed to hotcombustion gases and have generally linear geometries. For this reason,each vane airfoil portion 102 can be produced as a unitary piece from aceramic-based material, for example, a CMC material, and one or both ofthe platforms 104 can be produced from materials other than a ceramicmaterial. As represented in FIG. 23, each airfoil portion 102 comprisesa pair of oppositely-disposed nubs 106. Each nub 106 is effectively adovetail feature that defines an oppositely-disposed pair of tangs 108.Furthermore, the nubs 106 are entirely encased in the material used toform the platforms 104 so that the nubs 106 and their tangs 108 arereceived within pockets 110 that were defined in the platforms 104during a metal casting process used to form the platforms 104. Asevident from FIG. 23, each nub 106 and its tangs 108 define a region onthe airfoil portion 102 that is wider in cross-section than theimmediately adjacent region of the airfoil portion 102, and as such thenubs 106 are configured to serve as retention features for retaining theinner and outer platforms 104 on the airfoil portion 102, as well asretain the nubs 106 within the pockets 110 of the platforms 104. Otheraspects regarding the production of the vane assembly 100 can beappreciated from the discussions above regarding the turbine blades 30and 80 of FIGS. 3 through 21.

FIGS. 24 and 25 are isolated perspective views of the airfoil portions102 of FIG. 23 whose nubs 106 have been modified to include slots 112and holes 114, respectively, for promoting the attachment of the airfoilportions 102 to the metallic platforms 104. As with the slots 60 andholes 62 discussed in reference to FIGS. 11 and 12, during the processof casting the platforms 104 around the nubs 106, metallic materialenters the slots 112 and/or holes 114 such that solidification of thecasting material creates complementary interlocking metallic features(not shown) within the slots 112 and/or holes 114. In the case of holes114, the casting material within the holes 114 is also capable ofinterconnecting those portions 116 of the platforms 104 separated by thenubs 106. The interlocking effect physically promotes the retentioncapability provided by the nubs 106 and their tangs 108, and thereforefurther promotes a robust mechanical attachment between the airfoilportions 102 and the platforms 106. Though FIGS. 24 and 25 show theslots 112 and holes 114 as alternative configurations, combinations ofslots 112 and holes 114 are also within the scope of the invention.Furthermore, other negative surface features (depressions or recesses)could be defined in the nubs 106 to achieve a similar effect, andpositive surface features (not shown) could also be defined in the nubs106, similar to the protuberances 63 of FIG. 13.

While the invention has been described in terms of specific embodiments,it is apparent that other forms could be adopted by one skilled in theart. Therefore, the scope of the invention is to be limited only by thefollowing claims.

The invention claimed is:
 1. A process for producing a turbine bladecomprising a first subcomponent and at least a second subcomponent, thefirst subcomponent comprising an airfoil portion and a shank nub havinga base with a cross-section that is wider in cross-section than a regionof the shank nub adjacent the airfoil portion, the second subcomponentcomprising a dovetail portion attached to the shank nub and adapted foranchoring the turbine blade to a turbine disk, the dovetail portionhaving at least one off-axis geometric feature that results in thesecond subcomponent having a more complex geometry than the firstsubcomponent, the off-axis geometric feature comprising at least onetang that protrudes from the shank nub and a pressure face on the tang,the process comprising: producing the first subcomponent of aceramic-based material; and then casting a metallic material on theshank nub of the first subcomponent to form in situ the dovetail portionof the second subcomponent and the off-axis geometric feature thereofand to attach the second subcomponent to the first subcomponent, thesecond subcomponent being formed of at least the metallic material. 2.The process according to claim 1, wherein the casting of the metallicmaterial results in the metallic material entirely covering the base ofthe shank nub of the first subcomponent.
 3. The process according toclaim 1, wherein the ceramic-based material is chosen from the groupconsisting of homogeneous ceramic materials and ceramic matrix compositematerials containing a discontinuous and/or continuous ceramicreinforcement material in a ceramic matrix reinforcement material. 4.The process according to claim 1, wherein the ceramic-based material isa ceramic matrix composite material containing a continuous ceramicreinforcement material in a ceramic matrix reinforcement material, andat least one of the ceramic reinforcement material and the ceramicmatrix material comprises silicon carbide.
 5. The process according toclaim 1, wherein the ceramic-based material is a ceramic matrixcomposite material containing a discontinuous ceramic reinforcementmaterial in a ceramic matrix reinforcement material, and at least one ofthe ceramic reinforcement material and the ceramic matrix materialcomprises silicon carbide.
 6. The process according to claim 1, furthercomprising machining the metallic material to further define theoff-axis geometric feature.
 7. The process according to claim 1, whereinthe casting of the metal material results in the metal material notcovering any portion of the base of the shank nub and the base defines asecond tang that protrudes from the shank nub.
 8. The process accordingto claim 1, further comprising the step of applying a coating system onthe first subcomponent that inhibits chemical reactions between themetallic material of the second subcomponent and the ceramic-basedmaterial of the first subcomponent during the casting step.
 9. Theprocess according to claim 1, further comprising the step of defining atleast one recess in a surface region of the first subcomponent, and thecasting step results in some of the metallic material entering the atleast one recess to interlock the second subcomponent to the firstsubcomponent.
 10. The process according to claim 9, wherein the at leastone recess comprises at least one slot in the surface region of thefirst subcomponent, and the casting step results in some of the metallicmaterial entering the slot to interlock the second subcomponent to thefirst subcomponent.
 11. The process according to claim 1, wherein thesecond subcomponent further comprises a platform portion between theairfoil portion and the dovetail portion.
 12. The process according toclaim 1, wherein the first subcomponent further comprises a blade tipnub on the airfoil portion and the blade tip nub has a base with across-section that is wider than the airfoil portion, the method furthercomprising casting a metallic material around the blade tip nub of thefirst subcomponent to form in situ a shroud portion of the turbineblade, the shroud portion having an off-axis geometric featurecomprising at least one seal tooth that protrudes from the shroudportion.
 13. The turbine blade produced by the process of claim
 12. 14.The process according to claim 1, wherein the step of producing thefirst subcomponent comprises fabricating multiple prepregs that areindividually defined by a laminate comprising a fiber array in a matrixprecursor, stacking the multiple prepregs to form a laminate preform,and then firing the laminate preform to form the airfoil portion and theshank nub of the first subcomponent.
 15. The turbine blade produced bythe process of claim
 14. 16. The process according to claim 1, furthercomprising installing the turbine blade in a turbine section of a gasturbine engine by interlocking the dovetail portion with a complementarydovetail slot in a turbine disk.
 17. The turbine blade produced by theprocess of claim
 1. 18. A process for producing a component comprising afirst subcomponent and at least a second subcomponent having at leastone off-axis geometric feature that results in the second subcomponenthaving a more complex geometry than the first subcomponent, the processcomprising: producing the first subcomponent of a ceramic-basedmaterial; defining at least one recess in a surface region of the firstsubcomponent; and then separately forming the second subcomponent andthe off-axis geometric feature thereof and attaching the secondsubcomponent to the first subcomponent, the second subcomponent beingformed of at least one metallic material; wherein the steps ofseparately forming the second subcomponent and attaching the secondsubcomponent to the first subcomponent comprise casting the metallicmaterial around a region of the first subcomponent; wherein the castingof the metallic material around a region of the first subcomponentresults in some of the metallic material entering the at least onerecess to interlock the second subcomponent to the first subcomponent;and wherein the at least one recess comprises at least one hole throughthe first subcomponent, and the casting of the metallic material resultsin some of the metallic material entering the at least one hole tointerconnect regions of the second subcomponent that are separated bythe first subcomponent.
 19. A process for producing a componentcomprising a first subcomponent and at least a second subcomponenthaving at least one off-axis geometric feature that results in thesecond subcomponent having a more complex geometry than the firstsubcomponent, the process comprising: producing the first subcomponentof a ceramic-based material; and then separately forming the secondsubcomponent and the off-axis geometric feature thereof and attachingthe second subcomponent to the first subcomponent, the secondsubcomponent being formed of at least one metallic material; wherein thestep of separately forming the second subcomponent comprisesprefabricating the metallic material to form pieces of the secondsubcomponent, and the step of attaching the second subcomponent to thefirst subcomponent comprises assembling the pieces of the secondsubcomponent around a region of the first subcomponent, metallurgicallyjoining the pieces to each other, and then machining the pieces tocreate the second subcomponent and the off-axis geometric featurethereof.